The present invention relates generally to gas turbine engines, and, more specifically, to turbines therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine which powers the compressor, and additional energy is extracted in a low pressure turbine which typically powers an upstream fan in a typical turbofan aircraft engine application.
Engine efficiency and performance can be increased by increasing temperature of the combustion gases, but the hot combustion gases affect the life of turbine components heated thereby. Typical components such as nozzle vanes and rotor blades in the turbines are bathed with the hot combustion gases during operation and are typically cooled for prolonging their useful life in the engine.
For example, pressurized air is suitably bled from the compressor and channeled to the stationary nozzle vanes and rotating turbine rotor blades during operation for cooling thereof. The vanes and blades have correspondingly shaped hollow airfoils with internal cooling circuits therein.
Turbine airfoil cooling is quite esoteric and quite sophisticated, and the prior art is replete with a myriad of patents for maximizing cooling performance of the cooling air in the various regions of the airfoils.
The typical airfoil has a generally concave pressure side and generally convex suction side joined together at axially spaced apart leading and trailing edges. The airfoil extends from a radially inner root to a radially outer tip. For a turbine blade, the root is integral with a blade platform and the tip is spaced inwardly from a surrounding shroud. For a nozzle vane, both the root and the tip are integrally joined with corresponding inner and outer bands.
Inside the airfoil may be several cooling circuits with various configurations which typically discharge the cooling air through rows of film cooling holes in the pressure and suction sides of the airfoil. The cooling circuits include radial flow channels, some of which may be arranged end-to-end in serpentine fashion extending toward the leading edge of the airfoil or its trailing edge.
Small turbulator ribs may be formed on the inner surfaces of the airfoil for increasing heat transfer. Impingement cooling holes may be provided in corresponding bridges for impingement cooling the inner surface of the airfoil, typically at the hot leading edge. And, arrays of pins may be configured in two dimensional mesh grids for enhancing heat transfer cooling inside the airfoils.
Various patents in the prior art disclose typical embodiments of mesh cooling. Further advances in mesh cooling are presently being developed and are found, for example, in U.S. patent application Ser. No. 10/616,023 filed Jul. 9, 2003 now U.S. Pat. No. 6,832,889; Ser. No. 10/692,700 filed Oct. 24, 2003; and Ser. No. 10/718,465 filed Nov. 20, 2003, all assigned to the present assignee.
The various forms of cooling features in turbine airfoils are in many cases relatively small and must be capable of practical manufacture. For example, typical turbine airfoils are made by casting typical superalloy metals using corresponding ceramic cores which define the internal flow passages of the airfoil. The various impingement cooling holes, turbulator ribs, and mesh pins may be integrally formed in the cast airfoil by using corresponding features in the ceramic core or cores.
The individual radial passages in the turbine airfoil are formed by a corresponding ceramic core in the form of a slender leg or finger. The ceramic cores are relatively brittle and subject to damage during the casting process. If the cores are too thin or weak and prone to breakage, the effective yield of the casting process is reduced which correspondingly increases casting cost.
Following the casting process, the various rows of film cooling holes in their various simple to complex configurations may be formed using suitable drilling processes including laser drilling or electrical discharge machining (EDM) for example.
Since the leading edge region of the turbine airfoils first receives the hot combustion gases which flow thereover during operation, they are typically subject to the highest heat loads during operation and therefore require maximum cooling capability. Leading edge cooling configurations are myriad in the prior art.
Maximum leading edge cooling may be typically effected by providing impingement cooling directly behind the leading edge, and including several rows of showerhead film cooling holes through the leading edge for discharging the spent impingement cooling air in thermally insulating films over the external surface of the airfoil.
Aft of the leading edge in the suction side of the airfoil is typically found a row of film cooling gill holes for re-initiating the cooling film aft therefrom. And, aft of the leading edge on the pressure side are also found several rows of film cooling holes for re-initiating the cooling air films downstream therefrom.
In view of the typical complexity in effectively cooling the leading edge region of turbine airfoils, it is desired to provide an airfoil having improved leading edge cooling which may be introduced using conventional casting processes.